Method of upgrading a modular gas turbine engine

ABSTRACT

A method of upgrading a modular gas turbine engine, which includes: a first fan module with a fan having plurality of fan blades; a first engine core module including an engine core and a gearbox providing drive to the fan; and a first fan case module with a fan case arranged to enclose the fan blades, the method including: disassembling the gas turbine engine, replacing one of the first fan module, first engine core module or first fan case module with a replacement fan module, a replacement engine core module or a replacement fan case module; and reassembling the gas turbine engine using the replacement module, which is compatible with the others of the first fan module, first engine core module or first fan case module; and the replacement module designed to different parameters to one of the first fan module, first engine core module or first fan case module.

The present disclosure relates to a method of upgrading a modular gasturbine engine, and a system comprising a plurality of gas turbineengines.

Gas turbine aircraft engines comprise a propulsive fan arrangeddownstream of an air intake. The fan is surrounded by a fan case, andtypically generates two separate airflows. A first airflow is receivedby a core of the engine, and a second airflow is received in a bypassduct. The core comprises one or more compressors, a combustor, and oneor more turbines. The bypass duct is defined around the core.

In use, the core airflow is compressed by the compressors, mixed withfuel and combusted in the combustor. The combustion products areexpanded through the turbine stages and exhausted through a core nozzle.The turbines drive the compressor stages and propulsive fan through oneor more interconnecting shafts.

Typically, whilst some thrust is provided by the core nozzle, themajority of the thrust generated by the engine is provided by thepropulsive fan, through the bypass duct. Propulsive efficiency of thegas turbine can be improved by increasing the bypass ratio (the ratio ofthe air mass flow through the bypass duct to the air mass flow throughthe core). The bypass ratio is related to the size of the fan which inturn is limited by the rotation speed of the fan, as a large fanrotating at high speed may experience unwanted distortion of the fan,and other effects.

If the fan is driven by a reduction gearbox, it can be driven at slowerspeeds than the shafts from the turbines. This enables the fan to beincreased in size, facilitating an increase of the bypass duct ratio.

Generally, when an upgraded component is available for a gas turbineengine, the whole engine is taken out of use whilst the replacement partis fitted. Alternatively, if the upgrade is a significant part, a wholenew engine is provided.

According to a first aspect, there is provided a method of upgrading amodular gas turbine engine, wherein the gas turbine engine comprises: afirst fan module comprising a fan having plurality of fan blades; afirst engine core module including an engine core and a gearbox arrangedto provide drive to the fan; and a first fan case module comprising afan case arranged to enclose the fan blades, the method including thesteps of: disassembling the gas turbine engine, replacing a one of thefirst fan module, first engine core module or first fan case module witha replacement fan module, a replacement engine core module or areplacement fan case module; and reassembling the gas turbine engineusing the replacement module, wherein the replacement module iscompatible with the others of the first fan module, first engine coremodule or first fan case module; and wherein the replacement module isdesigned to different parameters to the one of the first fan module,first engine core module or first fan case module, such that it altersthe performance of the engine.

The method provides a way to upgrade an engine to provide improvedperformance, without having to take the whole engine out of use. Themethod therefore ensures that engines, and aircraft to which the enginesare secured, are not out of use for prolonged periods. This also allowsfor quick swapping of modules, and reduced development time through there-use of the modules.

The one of the first modules may be releasably secured to the others ofthe first modules through one or more connection points at one or moreconnection positions. The replacement module may be releasably securedto the others of the first modules through the one or more connectionpoints at the one or more connection positions.

The one of the first modules may include one or more apertures, openingsor electrical connections coupled to the others of the first modules infirst coupling positions. The replacement module may include the sameone or more apertures, openings or electrical connections coupled to theothers of the first modules in the first coupling positions.

The replacement module may be arranged to provide improved performanceof the gas turbine engine after reassembly. The improved performance mayinclude one or more of:

-   -   resistance to wear and tear or fatigue of one or more        components;    -   maximum thrust output;    -   engine efficiency;    -   engine noise;    -   engine weight;    -   engine emissions.

Replacing the one of the first modules may comprise: replacing one ormore component parts of a different fan module, engine core module orfan case module having the same design parameters as the one of thefirst modules, to form the replacement module.

The one of the first modules may comprise the fan module or the enginecore module. The method may further comprise: passing off thereassembled engine or the replacement module, to determine a powersetting parameter of the engine (10) or a power setting parameter orpower rating of the replacement module (50′, 52′, 54′) for use incontrolling the engine, during operation. Passing off only the upgradedengine or module allows the engine to be operated accurately andefficiently, in use.

Disassembling the gas turbine engine may comprise: removing the firstfan module from engine core; disengaging joints between the first enginecore module and the first fan case module; and separating the enginecore from the fan case in an axial direction.

The method may include replacing another of the fan module, engine coremodule or fan case module with a second replacement module, prior toreassembling, the second replacement module being designed to differentparameters to the other of the first fan module, first engine coremodule or first fan case module, such that it alters the performance ofthe engine.

The gearbox may be a first gearbox. The method may further include:replacing the first gearbox in the engine core module with a replacementgearbox prior to reassembling, the replacement gearbox being designed todifferent parameters to the first gearbox.

The engine core module and fan module may be passed off separately. Themethod may further comprise: passing off the replacement moduleindependently of the other of the first modules after reassembly of theengine. Passing off the fan and engine core separately ensures thatcomponents that are not replaced do not go through unnecessaryprocedures, and are not taken out of use unnecessarily.

According to a second aspect, there is provided a system comprising aplurality of gas turbine engines, the system having: a plurality offirst engine core modules, each comprising: an engine core having aturbine, a compressor, a core shaft connecting the turbine to thecompressor; and a gearbox that receives an input from the core shaft andoutputs drive at a lower rotational speed than the core shaft; aplurality of first fan modules compatible with the engine core firstmodules, each first fan module comprising a fan arranged to receivedrive from an engine core and having a plurality of fan blades; aplurality of first fan case modules compatible with the first enginecore modules and the first fan modules, each first fan case modulecomprising a fan case arranged to surround a fan and at least partiallydefining a bypass duct around an engine core; and a plurality ofreplacement modules comprising replacements for a one of the first fanmodules, the first engine core modules or the first fan case module,wherein the replacement modules are compatible with the others of thefirst fan modules, the first engine core modules or the first fan casemodules; wherein each of the plurality of gas turbine engines comprisesa first fan module from the plurality of first fan modules, a firstengine core module from the plurality of first engine core modules and afirst fan case module from the plurality of first fan case modules; andwherein the replacement modules are interchangeable with the one of thefirst fan modules, the first engine core modules or the first fan casemodule, such that it alters the performance of the engine.

The system provides a way to upgrade engines in a fleet, to provideimproved performance, without having to take the engines out of use.This is because the first modules can be swapped out for the replacementmodules. The system therefore ensures that engines, and aircraft towhich the engines are secured, are not out of use for prolonged periods.This also allows for quick swapping of modules, and reduced developmenttime through the re-use of the modules.

The one of the first modules may be releasably secured to the others ofthe first modules through one or more connection points at one or moreconnection positions. The replacement module may be releasably securedto the others of the first modules through the one or more connectionpoints at the one or more connection positions.

The one of the first modules may include one or more apertures, openingsor electrical connections coupled to the others of the first modules infirst coupling positions. The replacement module may include the sameone or more apertures, openings or electrical connections coupled to theothers of the first modules in the first coupling positions.

The replacement module may be arranged to provide improved performanceof the gas turbine engine after reassembly. The improved performance mayinclude one or more of:

-   -   resistance to wear and tear or fatigue of one or more        components;    -   maximum thrust output;    -   engine efficiency;    -   engine noise;    -   engine weight;    -   engine emissions.

The engine core module and fan module may be passed off separately. Themethod may further comprise: passing off the replacement moduleindependently of the other of the first modules after reassembly of theengine. Passing off the fan and engine core separately ensures thatcomponents that are not replaced do not go through unnecessaryprocedures, and are not taken out of use unnecessarily.

As noted elsewhere herein, the present disclosure may relate to a gasturbine engine. Such a gas turbine engine may comprise an engine corecomprising a turbine, a combustor, a compressor, and a core shaftconnecting the turbine to the compressor. Such a gas turbine engine maycomprise a fan (having fan blades) located upstream of the engine core.

Arrangements of the present disclosure may be particularly, although notexclusively, beneficial for fans that are driven via a gearbox.Accordingly, the gas turbine engine may comprise a gearbox that receivesan input from the core shaft and outputs drive to the fan so as to drivethe fan at a lower rotational speed than the core shaft. The input tothe gearbox may be directly from the core shaft, or indirectly from thecore shaft, for example via a spur shaft and/or gear. The core shaft mayrigidly connect the turbine and the compressor, such that the turbineand compressor rotate at the same speed (with the fan rotating at alower speed).

The gas turbine engine as described and/or claimed herein may have anysuitable general architecture. For example, the gas turbine engine mayhave any desired number of shafts that connect turbines and compressors,for example one, two or three shafts. Purely by way of example, theturbine connected to the core shaft may be a first turbine, thecompressor connected to the core shaft may be a first compressor, andthe core shaft may be a first core shaft. The engine core may furthercomprise a second turbine, a second compressor, and a second core shaftconnecting the second turbine to the second compressor. The secondturbine, second compressor, and second core shaft may be arranged torotate at a higher rotational speed than the first core shaft.

In such an arrangement, the second compressor may be positioned axiallydownstream of the first compressor. The second compressor may bearranged to receive (for example directly receive, for example via agenerally annular duct) flow from the first compressor.

The gearbox may be arranged to be driven by the core shaft that isconfigured to rotate (for example in use) at the lowest rotational speed(for example the first core shaft in the example above). For example,the gearbox may be arranged to be driven only by the core shaft that isconfigured to rotate (for example in use) at the lowest rotational speed(for example only be the first core shaft, and not the second coreshaft, in the example above). Alternatively, the gearbox may be arrangedto be driven by any one or more shafts, for example the first and/orsecond shafts in the example above.

In any gas turbine engine as described and/or claimed herein, acombustor may be provided axially downstream of the fan andcompressor(s). For example, the combustor may be directly downstream of(for example at the exit of) the second compressor, where a secondcompressor is provided. By way of further example, the flow at the exitto the combustor may be provided to the inlet of the second turbine,where a second turbine is provided. The combustor may be providedupstream of the turbine(s).

The or each compressor (for example the first compressor and secondcompressor as described above) may comprise any number of stages, forexample multiple stages. Each stage may comprise a row of rotor bladesand a row of stator vanes, which may be variable stator vanes (in thattheir angle of incidence may be variable). The row of rotor blades andthe row of stator vanes may be axially offset from each other.

The or each turbine (for example the first turbine and second turbine asdescribed above) may comprise any number of stages, for example multiplestages. Each stage may comprise a row of rotor blades and a row ofstator vanes. The row of rotor blades and the row of stator vanes may beaxially offset from each other.

Each fan blade may be defined as having a radial span extending from aroot (or hub) at a radially inner gas-washed location, or 0% spanposition, to a tip at a 100% span position. The ratio of the radius ofthe fan blade at the hub to the radius of the fan blade at the tip maybe less than (or on the order of) any of: 0.4, 0.39, 0.38 0.37, 0.36,0.35, 0.34, 0.33, 0.32, 0.31, 0.3, 0.29, 0.28, 0.27, 0.26, or 0.25. Theratio of the radius of the fan blade at the hub to the radius of the fanblade at the tip may be in an inclusive range bounded by any two of thevalues in the previous sentence (i.e. the values may form upper or lowerbounds). These ratios may commonly be referred to as the hub-to-tipratio. The radius at the hub and the radius at the tip may both bemeasured at the leading edge (or axially forwardmost) part of the blade.The hub-to-tip ratio refers, of course, to the gas-washed portion of thefan blade, i.e. the portion radially outside any platform.

The radius of the fan may be measured between the engine centreline andthe tip of a fan blade at its leading edge. The fan diameter (which maysimply be twice the radius of the fan) may be greater than (or on theorder of) any of: 250 cm (around 100 inches), 260 cm, 270 cm (around 105inches), 280 cm (around 110 inches), 290 cm (around 115 inches), 300 cm(around 120 inches), 310 cm, 320 cm (around 125 inches), 330 cm (around130 inches), 340 cm (around 135 inches), 350 cm, 360 cm (around 140inches), 370 cm (around 145 inches), 380 (around 150 inches) cm or 390cm (around 155 inches). The fan diameter may be in an inclusive rangebounded by any two of the values in the previous sentence (i.e. thevalues may form upper or lower bounds).

The rotational speed of the fan may vary in use. Generally, therotational speed is lower for fans with a higher diameter. Purely by wayof non-limitative example, the rotational speed of the fan at cruiseconditions may be less than 2500 rpm, for example less than 2300 rpm.Purely by way of further non-limitative example, the rotational speed ofthe fan at cruise conditions for an engine having a fan diameter in therange of from 250 cm to 300 cm (for example 250 cm to 280 cm) may be inthe range of from 1700 rpm to 2500 rpm, for example in the range of from1800 rpm to 2300 rpm, for example in the range of from 1900 rpm to 2100rpm. Purely by way of further non-limitative example, the rotationalspeed of the fan at cruise conditions for an engine having a fandiameter in the range of from 320 cm to 380 cm may be in the range offrom 1200 rpm to 2000 rpm, for example in the range of from 1300 rpm to1800 rpm, for example in the range of from 1400 rpm to 1600 rpm.

In use of the gas turbine engine, the fan (with associated fan blades)rotates about a rotational axis. This rotation results in the tip of thefan blade moving with a velocity U_(tip). The work done by the fanblades 13 on the flow results in an enthalpy rise dH of the flow. A fantip loading may be defined as dH/U_(tip) ², where dH is the enthalpyrise (for example the 1-D average enthalpy rise) across the fan andU_(tip) is the (translational) velocity of the fan tip, for example atthe leading edge of the tip (which may be defined as fan tip radius atleading edge multiplied by angular speed). The fan tip loading at cruiseconditions may be greater than (or on the order of) any of: 0.3, 0.31,0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39 or 0.4 (all units in thisparagraph being Jkg⁻¹K⁻¹/(ms⁻¹)²). The fan tip loading may be in aninclusive range bounded by any two of the values in the previoussentence (i.e. the values may form upper or lower bounds).

Gas turbine engines in accordance with the present disclosure may haveany desired bypass ratio, where the bypass ratio is defined as the ratioof the mass flow rate of the flow through the bypass duct to the massflow rate of the flow through the core at cruise conditions. In somearrangements the bypass ratio may be greater than (or on the order of)any of the following: 10, 10.5, 11, 11.5, 12, 12.5, 13, 13.5, 14, 14.5,15, 15.5, 16, 16.5, or 17. The bypass ratio may be in an inclusive rangebounded by any two of the values in the previous sentence (i.e. thevalues may form upper or lower bounds). The bypass duct may besubstantially annular. The bypass duct may be radially outside the coreengine. The radially outer surface of the bypass duct may be defined bya nacelle and/or a fan case.

The overall pressure ratio of a gas turbine engine as described and/orclaimed herein may be defined as the ratio of the stagnation pressureupstream of the fan to the stagnation pressure at the exit of thehighest pressure compressor (before entry into the combustor). By way ofnon-limitative example, the overall pressure ratio of a gas turbineengine as described and/or claimed herein at cruise may be greater than(or on the order of) any of the following: 35, 40, 45, 50, 55, 60, 65,70, 75. The overall pressure ratio may be in an inclusive range boundedby any two of the values in the previous sentence (i.e. the values mayform upper or lower bounds).

Specific thrust of an engine may be defined as the net thrust of theengine divided by the total mass flow through the engine. At cruiseconditions, the specific thrust of an engine described and/or claimedherein may be less than (or on the order of) any of the following: 110Nkg⁻¹'s, 105 Nkg⁻¹ s, 100 Nkg⁻¹ s, 95 Nkg⁻¹ s, 90 Nkg⁻¹ s, 85 Nkg⁻¹ s or80 Nkg⁻¹ s. The specific thrust may be in an inclusive range bounded byany two of the values in the previous sentence (i.e. the values may formupper or lower bounds). Such engines may be particularly efficient incomparison with conventional gas turbine engines.

A gas turbine engine as described and/or claimed herein may have anydesired maximum thrust. Purely by way of non-limitative example, a gasturbine as described and/or claimed herein may be capable of producing amaximum thrust of at least (or on the order of) any of the following:160 kN, 170 kN, 180 kN, 190 kN, 200 kN, 250 kN, 300 kN, 350 kN, 400 kN,450 kN, 500 kN, or 550 kN. The maximum thrust may be in an inclusiverange bounded by any two of the values in the previous sentence (i.e.the values may form upper or lower bounds). The thrust referred to abovemay be the maximum net thrust at standard atmospheric conditions at sealevel plus 15 deg C. (ambient pressure 101.3 kPa, temperature 30 degC.), with the engine static.

In use, the temperature of the flow at the entry to the high pressureturbine may be particularly high. This temperature, which may bereferred to as TET, may be measured at the exit to the combustor, forexample immediately upstream of the first turbine vane, which itself maybe referred to as a nozzle guide vane. At cruise, the TET may be atleast (or on the order of) any of the following: 1400K, 1450K, 1500K,1550K, 1600K or 1650K. The TET at cruise may be in an inclusive rangebounded by any two of the values in the previous sentence (i.e. thevalues may form upper or lower bounds). The maximum TET in use of theengine may be, for example, at least (or on the order of) any of thefollowing: 1700K, 1750K, 1800K, 1850K, 1900K, 1950K or 2000K. Themaximum TET may be in an inclusive range bounded by any two of thevalues in the previous sentence (i.e. the values may form upper or lowerbounds). The maximum TET may occur, for example, at a high thrustcondition, for example at a maximum take-off (MTO) condition.

A fan blade and/or aerofoil portion of a fan blade described and/orclaimed herein may be manufactured from any suitable material orcombination of materials. For example at least a part of the fan bladeand/or aerofoil may be manufactured at least in part from a composite,for example a metal matrix composite and/or an organic matrix composite,such as carbon fibre. By way of further example at least a part of thefan blade and/or aerofoil may be manufactured at least in part from ametal, such as a titanium based metal or an aluminium based material(such as an aluminium-lithium alloy) or a steel based material. The fanblade may comprise at least two regions manufactured using differentmaterials. For example, the fan blade may have a protective leadingedge, which may be manufactured using a material that is better able toresist impact (for example from birds, ice or other material) than therest of the blade. Such a leading edge may, for example, be manufacturedusing titanium or a titanium-based alloy. Thus, purely by way ofexample, the fan blade may have a carbon-fibre or aluminium based body(such as an aluminium lithium alloy) with a titanium leading edge.

A fan as described and/or claimed herein may comprise a central portion,from which the fan blades may extend, for example in a radial direction.The fan blades may be attached to the central portion in any desiredmanner. For example, each fan blade may comprise a fixture which mayengage a corresponding slot in the hub (or disc). Purely by way ofexample, such a fixture may be in the form of a dovetail that may slotinto and/or engage a corresponding slot in the hub/disc in order to fixthe fan blade to the hub/disc. By way of further example, the fan bladesmaybe formed integrally with a central portion. Such an arrangement maybe referred to as a blisk or a bling. Any suitable method may be used tomanufacture such a blisk or bling. For example, at least a part of thefan blades may be machined from a block and/or at least part of the fanblades may be attached to the hub/disc by welding, such as linearfriction welding.

The gas turbine engines described and/or claimed herein may or may notbe provided with a variable area nozzle (VAN). Such a variable areanozzle may allow the exit area of the bypass duct to be varied in use.The general principles of the present disclosure may apply to engineswith or without a VAN.

The fan of a gas turbine as described and/or claimed herein may have anydesired number of fan blades, for example 16, 18, 20, or 22 fan blades.

As used herein, cruise conditions may mean cruise conditions of anaircraft to which the gas turbine engine is attached. Such cruiseconditions may be conventionally defined as the conditions atmid-cruise, for example the conditions experienced by the aircraftand/or engine at the midpoint (in terms of time and/or distance) betweentop of climb and start of decent.

Purely by way of example, the forward speed at the cruise condition maybe any point in the range of from Mach 0.7 to 0.9, for example 0.75 to0.85, for example 0.76 to 0.84, for example 0.77 to 0.83, for example0.78 to 0.82, for example 0.79 to 0.81, for example on the order of Mach0.8, on the order of Mach 0.85 or in the range of from 0.8 to 0.85. Anysingle speed within these ranges may be the cruise condition. For someaircraft, the cruise conditions may be outside these ranges, for examplebelow Mach 0.7 or above Mach 0.9.

Purely by way of example, the cruise conditions may correspond tostandard atmospheric conditions at an altitude that is in the range offrom 10000 m to 15000 m, for example in the range of from 10000 m to12000 m, for example in the range of from 10400 m to 11600 m (around38000 ft), for example in the range of from 10500 m to 11500 m, forexample in the range of from 10600 m to 11400 m, for example in therange of from 10700 m (around 35000 ft) to 11300 m, for example in therange of from 10800 m to 11200 m, for example in the range of from 10900m to 11100 m, for example on the order of 11000 m. The cruise conditionsmay correspond to standard atmospheric conditions at any given altitudein these ranges.

Purely by way of example, the cruise conditions may correspond to: aforward Mach number of 0.8; a pressure of 23000 Pa; and a temperature of−55 deg C.

As used anywhere herein, “cruise” or “cruise conditions” may mean theaerodynamic design point. Such an aerodynamic design point (or ADP) maycorrespond to the conditions (comprising, for example, one or more ofthe Mach Number, environmental conditions and thrust requirement) forwhich the fan is designed to operate. This may mean, for example, theconditions at which the fan (or gas turbine engine) is designed to haveoptimum efficiency.

In use, a gas turbine engine described and/or claimed herein may operateat the cruise conditions defined elsewhere herein. Such cruiseconditions may be determined by the cruise conditions (for example themid-cruise conditions) of an aircraft to which at least one (for example2 or 4) gas turbine engine may be mounted in order to provide propulsivethrust.

The skilled person will appreciate that except where mutually exclusive,a feature or parameter described in relation to any one of the aboveaspects may be applied to any other aspect. Furthermore, except wheremutually exclusive, any feature or parameter described herein may beapplied to any aspect and/or combined with any other feature orparameter described herein.

Embodiments will now be described by way of example only, with referenceto the Figures, in which:

FIG. 1 is a sectional side view of a gas turbine engine;

FIG. 2 is a close up sectional side view of an upstream portion of a gasturbine engine;

FIG. 3 is a partially cut-away view of a gearbox for a gas turbineengine;

FIG. 4A illustrates a schematic view of the gas turbine engine of FIG.1, illustrating the separate modules of the engine;

FIG. 4B illustrates the modules of FIG. 4A, in exploded form;

FIG. 4C schematically illustrates a cut-through view of the gas turbineengine of FIGS. 4A and 4B, from the front, in the region of the outletguide vanes;

FIG. 5 schematically illustrates separation of an engine core module anda fan case module; and

FIG. 6 illustrates a method for replacing a module of the gas turbineengine.

FIG. 1 illustrates a gas turbine engine 10 having a principal rotationalaxis 9. The engine 10 comprises an air intake 12 and a propulsive fan 23that generates two airflows: a core airflow A and a bypass airflow B.The gas turbine engine 10 comprises a core 11 that receives the coreairflow A. The engine core 11 comprises, in axial flow series, a lowpressure compressor 14, a high-pressure compressor 15, combustionequipment 16, a high-pressure turbine 17, a low pressure turbine 19 anda core exhaust nozzle 20. The fan 23 is attached to and driven by thelow pressure turbine 19 via a shaft 26 and an epicyclic gearbox 30.

The propulsive fan 23 includes a plurality of fan blades 25 extendingradially outward from a hub 29 mounted on an output shaft of the gearbox30. The radially outer tips of the fan blades 25 are surrounded by a fancasing 42, which extends downstream behind the fan 23. Behind the fancasing 42, in the axial flow direction (downstream), a nacelle 21surrounds the engine core 11. The fan casing 42 and nacelle 21 define abypass duct 22 and a bypass exhaust nozzle 18 around the engine core 11.

The bypass airflow B flows through the bypass duct 22. At an upstreamend of the bypass duct 22, adjacent an intake 31 of the bypass duct 22,and downstream of the fan 23, a plurality of outlet guide vanes 33extend radially between the engine core 11 and the fan casing 42. Theoutlet guide vanes 33 reduce swirl and turbulence in the bypass airflowB, providing improved thrust.

In use, the core airflow A enters the core intake 35, and is acceleratedand compressed by the low pressure compressor 14 and directed into thehigh pressure compressor 15 where further compression takes place. Thecompressed air exhausted from the high pressure compressor 15 isdirected into the combustion equipment 16 where it is mixed with fueland the mixture is combusted. The resultant hot combustion products thenexpand through, and thereby drive, the high pressure and low pressureturbines 17, 19 before being exhausted through the nozzle 20 to providesome propulsive thrust. The high pressure turbine 17 drives the highpressure compressor 15 by a suitable interconnecting shaft 27. The fan23 generally provides the majority of the propulsive thrust. Theepicyclic gearbox 30 is a reduction gearbox.

An exemplary arrangement for a geared fan gas turbine engine 10 is shownin FIG. 2. The low pressure turbine 19 (see FIG. 1) drives the shaft 26,which is coupled to a sun wheel, or sun gear, 28 of the epicyclic geararrangement 30. Radially outwardly of the sun gear 28 and intermeshingtherewith is a plurality of planet gears 32 that are coupled together bya planet carrier 34. The planet carrier 34 constrains the planet gears32 to precess around the sun gear 28 in synchronicity whilst enablingeach planet gear 32 to rotate about its own axis. The planet carrier 34is coupled via linkages 36 to the fan 23 in order to drive its rotationabout the engine axis 9. Radially outwardly of the planet gears 32 andintermeshing therewith is an annulus or ring gear 38 that is coupled,via linkages 40, to a stationary supporting structure 24.

Note that the terms “low pressure turbine” and “low pressure compressor”as used herein may be taken to mean the lowest pressure turbine stagesand lowest pressure compressor stages (i.e. not including the fan 23)respectively and/or the turbine and compressor stages that are connectedtogether by the interconnecting shaft 26 with the lowest rotationalspeed in the engine (i.e. not including the gearbox output shaft thatdrives the fan 23). In some literature, the “low pressure turbine” and“low pressure compressor” referred to herein may alternatively be knownas the “intermediate pressure turbine” and “intermediate pressurecompressor”. Where such alternative nomenclature is used, the fan 23 maybe referred to as a first, or lowest pressure, compression stage.

The epicyclic gearbox 30 is shown by way of example in greater detail inFIG. 3. Each of the sun gear 28, planet gears 32 and ring gear 38comprise teeth about their periphery to intermesh with the other gears.However, for clarity only exemplary portions of the teeth areillustrated in FIG. 3. There are four planet gears 32 illustrated,although it will be apparent to the skilled reader that more or fewerplanet gears 32 may be provided within the scope of the claimedinvention. Practical applications of a planetary epicyclic gearbox 30generally comprise at least three planet gears 32.

The epicyclic gearbox 30 illustrated by way of example in FIGS. 2 and 3is of the planetary type, in that the planet carrier 34 is coupled to anoutput shaft via linkages 36, with the ring gear 38 fixed. However, anyother suitable type of epicyclic gearbox 30 may be used. By way offurther example, the epicyclic gearbox 30 may be a star arrangement, inwhich the planet carrier 34 is held fixed, with the ring (or annulus)gear 38 allowed to rotate. In such an arrangement the fan 23 is drivenby the ring gear 38. By way of further alternative example, the gearbox30 may be a differential gearbox in which the ring gear 38 and theplanet carrier 34 are both allowed to rotate.

It will be appreciated that the arrangement shown in FIGS. 2 and 3 is byway of example only, and various alternatives are within the scope ofthe present disclosure. Purely by way of example, any suitablearrangement may be used for locating the gearbox 30 in the engine 10and/or for connecting the gearbox 30 to the engine 10. By way of furtherexample, the connections (such as the linkages 36, 40 in the FIG. 2example) between the gearbox 30 and other parts of the engine 10 (suchas the input shaft 26, the output shaft and the fixed structure 24) mayhave any desired degree of stiffness or flexibility. By way of furtherexample, any suitable arrangement of the bearings between rotating andstationary parts of the engine (for example between the input and outputshafts from the gearbox and the fixed structures, such as the gearboxcasing) may be used, and the disclosure is not limited to the exemplaryarrangement of FIG. 2. For example, where the gearbox 30 has a stararrangement (described above), the skilled person would readilyunderstand that the arrangement of output and support linkages andbearing locations would typically be different to that shown by way ofexample in FIG. 2.

Accordingly, the present disclosure extends to a gas turbine enginehaving any arrangement of gearbox styles (for example star orplanetary), support structures, input and output shaft arrangement, andbearing locations.

Optionally, the gearbox may drive additional and/or alternativecomponents (e.g. the intermediate pressure compressor and/or a boostercompressor).

Other gas turbine engines to which the present disclosure may be appliedmay have alternative configurations. For example, such engines may havean alternative number of compressors and/or turbines and/or analternative number of interconnecting shafts. By way of further example,the gas turbine engine shown in FIG. 1 has a split flow nozzle 20, 22meaning that the flow through the bypass duct 22 has its own nozzle thatis separate to and radially outside the core engine nozzle 20. However,this is not limiting, and any aspect of the present disclosure may alsoapply to engines in which the flow through the bypass duct 22 and theflow through the core 11 are mixed, or combined, before (or upstream of)a single nozzle, which may be referred to as a mixed flow nozzle. One orboth nozzles (whether mixed or split flow) may have a fixed or variablearea. Whilst the described example relates to a turbofan engine, thedisclosure may apply, for example, to any type of gas turbine engine,such as an open rotor (in which the fan stage is not surrounded by anacelle) or turboprop engine, for example. The gas turbine engine 10 mayalso be arranged in the “pusher” configuration, in which the fan 23 islocated downstream of the core 11. In some arrangements, the gas turbineengine 10 may not comprise a gearbox 30.

The geometry of the gas turbine engine 10, and components thereof, isdefined by a conventional axis system, comprising an axial direction(which is aligned with the rotational axis 9), a radial direction (inthe bottom-to-top direction in FIG. 1), and a circumferential direction(perpendicular to the page in the FIG. 1 view). The axial, radial andcircumferential directions are mutually perpendicular.

FIG. 4A schematically illustrates the constituent components of the gasturbine engine 10 of FIGS. 1 to 3, with the nacelle 21 removed. As shownin FIG. 4B, the gas turbine engine 10 is formed of a number of separatemodules 50, 52, 54. The engine 10 may thus be considered modular.

The first module is an engine core module 52. This typically includesthe gearbox 30, low pressure compressor 14, high-pressure compressor 15,combustion equipment 16, high-pressure turbine 17, and low pressureturbine 19. The engine core module 52 can also be referred to as apropulsor.

The second module, also referred to as the fan module 50, includes thefan blades 25.

The third module is a fan case module 54. This includes the fan case 42,with the outlet guide vanes 33 extending inwardly from the fan case 42.The hub 29 and gearbox 30 may be part of the fan module 520 or theengine core module 52. The gearbox 30 may additionally be configured asa separable module in its own right or part of the fan case module 54.

In the assembled engine, the engine core module 52 is joined to the fancase module 54 by joints 56 formed at the radially inner ends 58 of theoutlet guide vanes 33. Slots 60 are formed in the outer surface 62 ofthe engine core 11. The slots 60 have a closed end 64 facing downstreamthrough the engine 10, and an open end 66 opposite the closed end 64.

A projection 68 is formed on the radially inner end 58 of each outletguide vane 33. The projection 68 is arranged to co-operate with the slot60 on the engine core 11. Once the projection 68 is received in the slot60, a closing member 72, such as a bolt, may be used to close the openend 66 of the slot 60.

The closing member 72, and closed end 64 prevent relative axial movementof the fan case module 54 and the engine core module 52. In one example,the closing member 72 may pass through a corresponding opening (notshown) formed in the projection to further prevent relative movement ofthe two modules 52, 54. The slot 60 and projection 68 may optionally beshaped to prevent radial movement of the outlet guide vanes 33, and thusfan casing 42, relative to the engine core 11.

FIG. 4C schematically illustrates a sectional front view of the engine10, facing the upstream end of the outlet guide vanes 33 of the engine.As can be seen in FIG. 4C, the closing members 72 are arranged aroundthe inner circumference of the bypass duct 22. As shown in FIG. 4C, thejoints 56 are thus formed around this circumference.

Further connecting/support struts (not shown) may also be providedbetween the fan case 42 and the engine core 11, if required.

The fan blades 25 of the fan module 50 may be secured to the output ofthe gearbox in any suitable manner. For example, the fan hub 29 and fanblades 25 may also have corresponding projections and slots (not shown)to fix the blades to the engine core 11. A nose cone 37 is also providedon the axially front end of the engine core 11, to retain the fan 23.

The fan case module 54 is secured to the nacelle 21 of the engine 10through suitable fixings. The engine 10 is then secured to an aircraft(not shown) through struts extending from the nacelle 21.

As shown in FIG. 4B, the fan module 50 can be separated from the enginecore module 52, and the engine core module 52 and fan case module 54 canbe separated from one another. The separation of the modules 50, 52, 54will now be discussed in more detail, with reference to FIG. 5.

As a first step of separating the modules 50, 52, 54, the fan module 50is removed from the engine core module 52. Prior to the removal of thefan blades 25, it may be necessary to remove the nose cone 37. Once thisis done, the fan blades may then be removed. FIG. 5 illustrates theengine 10 with the fan module 50 removed.

As discussed above, support struts (not shown) may also be providedbetween the fan case 42 and the engine core 11. Where such struts arepresent, these are disconnected. This may be before or after removal ofthe fan module 50.

Following this, the joints 56 formed between the outlet guide vanes 33and engine core 11 are disengaged by removing the closing members fromthe slots 60. These joints 56 provide the only structural link betweenthe engine core module 52 and the fan case module 54. Therefore, priorto removal, a lifting tool may be brought into support the engine coremodule 52, such that when joints 56 are disengaged, the core 11 is stillsupported.

The engine core 11 includes a number of lifting points 74 (see FIG. 5)at which the lifting tool can engage the core 11. The lifting points 74may be formed in the housing 76 of the core 11, and connected through tothe support structure 24 of the engine 10.

As shown in FIG. 4C, the joints 56 are arranged around the innercircumference of the bypass duct 22. In one example, the joints may bedisengaged in diametrically opposed pairs, until a single pair remains,to ensure that the load of the engine core 11 is always as evenly spreadas possible.

Once all joints have been disconnected, the lifting tool is used toextract the engine core module 52 from the fan case module 54. The coreis moved axially with respect to the principal axis 9′ of the engine 10,in the direction away from the closed ends 64 of the slots 60, shown bythe arrows 80.

It will be appreciated that there may be very little clearance betweenthe engine core module 52 and fan case module 54. This means that thetwo modules 52, 54 need to be aligned to a high degree of accuracy inorder to move the engine core module 52 out of the fan case module 54.Furthermore, as the joints 56 are disengaged, the engine core module 52may shift vertically with respect to the fan module 54 as the loaddistribution of the engine core 11 changes. To accommodate this change,the lifting tool may vertically adjust the height of the engine coremodule 52 during the process of separating the modules.

The fan case module 54 may also be disconnected from the nacelle 21. Aseparate lifting tool may be provided for the fan case module 54. Aswith the engine core module 52, the fan case module 54 may includecorresponding lifting points 78 formed in the fan case 42 and/or outletguide vanes 33. The relative vertical movement between the fan casemodule 54 and engine core module 52 may be provided by either the fancase lifting tool on its own, the propulsor lifting tool on its own, orboth in combination.

It will be appreciated that in some instances, the above process may becarried out with the engine 10 mounted on the wing of an aircraft. Inalternative examples, the engine may be removed from the wing, andsuspended in a support structure. The lifting tools for the engine coremodule 52 and/or fan case module 54 may be part of the supportstructure, or may be separate.

It will also be appreciated that the engine 10 may be assembled usingthe reverse of the above process.

The ability to separate the engine 10 into separate modules 50, 52, 54facilitates easy delivery and transport of the engine 10. Furthermore,the fan case 42 is typically continuous around its circumferences. Thismeans that the fan case module 54 is large and difficult to transport.However, the fan case module 54 requires significantly less service,repair and maintenance than the other modules 50, 52.

Therefore, the ability to remove the fan module 50 and the engine coremodule 52 means that these parts, which require more regular servicingbut which are easier to transport, can be shipped for servicing withouthaving to ship the fan case 42.

Gas turbine engines 10 typically include accessory drive units 82. Theseare arranged to take drive from the core shafts 26, 27 of the engine 10,and use it to power subsystems of the engine 10 and aircraft, such ascooling systems, cabin air systems and the like. In some examples of thegas turbine engine 10 discussed above, the accessory drive unit(s) 82 ofthe engine are provided within the engine core 11. Therefore, theaccessory unit(s) 82 are part of the engine core module 52.

Using core mounted accessory unit(s) 82 ensures that no drive needs tobe transmitted across the bypass duct 22. Furthermore, no cooling fluidor air for the accessory unit(s) 82 needs to be transmitted across thebypass duct 22 either. This simplifies the connection between themodules 50, 52, 54 since there is reduced connections between thedifferent modules 50, 52, 54. In some examples, there may be no air orfluid connections across the bypass duct 22 at all. Optionally,electrical connections may be provided across the bypass duct 22, but insome examples there may not even be electrical connections.

Examples of core mounted accessory units may include units such as theoil tank, accessory gearbox and related systems, data entry plug,ignitors, oil heat management systems, and the associated pipes andcables.

The different modules 50, 52, 54 are interchangeable. This means that,for example, a particular fan module 50 and fan case module 54 may beused with any engine core module 52, where the engine core module 54 ismade to the same configuration and design parameters.

Similarly, the engine core module 52 may be used with any fan module 50and fan case module 54, and any compatible fan case module 54 may beused with any compatible fan module 50.

The interchangeability of modules 50, 52, 54 means that, for example, afirst engine core module 52 may be swapped out for a different, secondengine core module 52′. In order for two different engine core modules52, 52′ to be interchangeable, there should, if possible, be a number ofcommon features between the modules 52, 52′.

In particular, the second engine core module 52′ should, if possible,connect to the fan case module 54 and the fan module 50 by joints 56that are the same as with the first engine core module 52. Theconfiguration and positions of the joints 56 should be the same. Anyother connection points, for example for struts and the like, shouldalso be provided in the same place and style.

Furthermore, where there are other links between the modules 50, 52, 54,such as conduits for air or fluid, electrical connections, dataconnections or other connections, these links are also provided in thesame locations on the first and second engine core modules 52, 52′.

Notwithstanding the above common features between the first and secondengine core modules 52, 52′, the second engine core module 52′ is madeto a different design to the first engine core module 52. As such,replacing the engine core module 52 can improve the engine performance.Thus, by replacing the engine core module 52, the whole engine 10 can beupgraded without having to take the aircraft fitted within the engine 10out of service for a prolonged period, and without having to provide newfan and fan case modules 50, 54.

The improved performance may be in one or more of a number of differentcategories. For example, the improved performance may relate to:resistance to wear and tear or fatigue of one or more components;maximum thrust output; engine efficiency; engine noise; engine weight;and engine emissions.

The design of the second engine core module 52′ may vary from the firstengine core module 52 in a number of different ways. This may be in thematerials used for some components, the size, shape or relative locationof components, new schematic layouts, or any other design parameter ofthe engine 10.

FIG. 6 illustrates a method 600 in which one of the modules 50, 52, 54in an engine 10, such as the engine core module 52, can be replaced andupgraded.

In a first step 602, the engine 10 is disassembled. This may includeremoving the fan module 50 from the engine core module 52, and removingthe engine core 52 from the fan casing module 54, as discussed above.

In a second step 604, the engine core module 52 is switched with adifferent engine core module 52′, compatible with the engine 10. In athird step 606, the engine 10 is reassembled using the original fanmodule 50 and fan case module 54, and the replacement engine core module52′. In this way, the aircraft to which the engine 10 is fitted can bekept in use, whilst the engine core module 52 is serviced.

There are two contributions to the thrust generated by a gas turbineengine 10. The first is the core airflow B through the engine core 11,the second is the bypass airflow A, through the bypass duct 22. Inflight, the thrust produced by a gas turbine aircraft engine 10 cannotbe directly measured. In one method of operating a gas turbine engine10, a power setting parameter is used to determine the thrust producedby the engine 10. The power setting parameter converts a measureablevariable of the engine (such as shaft rotation speed) to the totalthrust, to enable control of the engine 10.

The power setting parameter is established by a process known as passingoff. In passing off, the engine 10 is calibrated in a test rig able tomeasure the thrust to determine the relationship between the thrust andthe measureable variable.

Once the engine 10 is reassembled, according to the above method, theengine 10 can be passed off, to determine the new power settingparameter. Alternatively, the separate modules 50, 52, 54 of the enginecan be passed off separately, to provide a separate power rating foreach module 50, 52, 54. Therefore, where a module 50, 52, 54 isreplaced, the reassembled engine 10 does not require passing off.Instead, the power setting parameter will combine the power ratings forthe set of modules 50, 52, 54, 50′, 52′, 54′ combined to form the engine10, including any original modules 50, 52, 54 and any upgraded modules50′, 52′, 54′.

In other examples, each module 50, 52, 54 may have a separate powersetting parameter that enables the contribution of each module to thetotal thrust to be determined. Again, when a module 50, 52, 54 isreplaced, these contributions can then be totalled to determine thetotal thrust generated by the engine 10, without having to pass off thewhole engine.

In the examples where the engine does not require passing off as acompleted unit when a module 50, 52, 54 is upgraded, only the new module50′, 52′ needs passing off. The modules 50, 52, 54 that are not replacedare not passed off.

An operator of a fleet of aircraft or engines 10 may have a plurality offirst engine core modules 52, a first plurality of fan modules 50 and afirst plurality of fan case modules 54. The operator may also have aplurality of second engine core modules 52′, and/or a second pluralityof fan modules 50′ and/or a second plurality of fan case modules 54′. Anengine 10 can be upgraded by replacement of one of the first module 50,52, 54 with one of the second modules 50′, 52′, 54′. Each engine 10 mayinitially include any one of each of the first engine core modules 52,any one of the first fan modules 50 and any one of the first fan casemodules 54, rather than each engine 10 comprising dedicated sets ofmodules 50, 52, 54 that can only be used together. Theinterchangeability of modules 50, 52, 54 allows the engine core modules52 (or other modules 50, 52) to be serviced, replaced, repaired orupgraded, whilst aircraft fitted the engines 10 remain functional.Otherwise, a replacement engine 10 would need to be transported to theaircraft, or the aircraft would have to be taken out of service for theengines 10 to be serviced.

The system may include more of the different modules 50, 52, 54, thanengines 10. This ensures that aircraft can be kept in service whilst themodules undergo maintenance. In particular, but not exclusively, theremay be more engine core modules 52 than engines 10, as this partrequires the most regular maintenance.

It will be appreciated that the joints 56 between the outlet guide vanes33 and the engine core 11 discussed above are just one example way ofconnecting the fan case module 54 to the engine core module 52. In somecases, the slots 60 may be open at both ends, with respective closingmembers 70, or the slot 60 may be open at either end.

The slot may be formed in the housing 76 of the engine core 11, with athrough connection to the support structure 24 of the engine 10, or maybe formed and coupled to the support structure 24 in any other way.

Any type of joint may be used between the outlet guide vanes 33 and theengine core 11, instead of the joint using slots 60 and projections 68.In one alternative example, the engine core 11 may include a pair ofradial extending flanges (not shown) positioned at the upstream anddownstream ends of the outlet guide vanes 33. The flanges are used tobolt the guide vanes 33 to the engine core 11. The flanges may becontinuous around the outside of the engine core 11, or may bediscontinuous. Where the flanges are discontinuous, sufficient flangesmay be provided to couple all of the outlet guide vanes 33 to the enginecore 11.

Furthermore, the provision of joints 56 at this particular location isgiven by way of example, only. The joints may be provided in anysuitable location in the engine 10.

The method of separating the modules 50, 52, 54 discussed above is givenby way of example only. Any suitable method may be used to separate themodules 50, 52, 54. Furthermore, any of the modules 50, 52, 54 may beupgraded, instead of the engine core module 52. In some cases, two ofthe modules 50, 52, 54 may be upgraded at the same time. In someexamples, the gearbox 30 may be considered to be a further upgradeablemodule, in addition to the fan module 50, engine core module 52, and fancase module 54.

In some cases, the replacement modules 50′, 52′, 54′ may be newlymanufactured modules. However, it will be appreciated that in othercases, where an operator has a plurality of each of the modules 50, 52,54, existing modules 50, 52, 54 from the plurality may be overhauled andmodified to form the replacement modules 50′, 52′, 54′.

In some cases, the joints 56, links and connections on the replacementmodule 50′, 52′, 54′ may be in different, but similar locations to thelinks on the first module 50, 52, 54, or of different format, ordifferent in other ways. It will be appreciated that in some cases, thelinks and connections may be able to accommodate small differences. Inother cases, adapters or converters may be provided to accommodatedifferences.

It will be appreciated that in some cases, the replacement module 50′,52′, 54′ may downgrade engine performance in some of the areas discussedabove. For example, an engine core module 52 may be replaced with amodule generating lower thrust where high levels of thrust are notneeded, and operation of a high thrust engine would be inefficient.

It will be understood that the invention is not limited to theembodiments above-described and various modifications and improvementscan be made without departing from the concepts described herein. Exceptwhere mutually exclusive, any of the features may be employed separatelyor in combination with any other features and the disclosure extends toand includes all combinations and sub-combinations of one or morefeatures described herein.

1. A method of upgrading a modular gas turbine engine, wherein the gasturbine engine comprises: a first fan module comprising a fan havingplurality of fan blades; a first engine core module including an enginecore and a gearbox arranged to provide drive to the fan; and a first fancase module comprising a fan case arranged to enclose the fan blades,the method including the steps of: disassembling the gas turbine engine,replacing a one of the first fan module, first engine core module orfirst fan case module with a replacement fan module, a replacementengine core module or a replacement fan case module; and reassemblingthe gas turbine engine using the replacement module, wherein thereplacement module is compatible with the others of the first fanmodule, first engine core module or first fan case module; and whereinthe replacement module is designed to different parameters to the one ofthe first fan module, first engine core module or first fan case module,such that it alters the performance of the engine.
 2. The method ofclaim 1, wherein: the one of the first modules is releasably secured tothe others of the first modules through one or more connection points atone or more connection positions; and the replacement module isreleasably secured to the others of the first modules through the one ormore connection points at the one or more connection positions.
 3. Themethod of claim 1, wherein: the one of the first modules includes one ormore apertures, openings or electrical connections coupled to the othersof the first modules in first coupling positions; and the replacementmodule includes the same one or more apertures, openings or electricalconnections coupled to the others of the first modules in the firstcoupling positions.
 4. The method of claim 1, wherein the replacementmodule is arranged to provide improved performance of the gas turbineengine after reassembly.
 5. The method of claim 4, wherein the improvedperformance may include one or more of: resistance to wear and tear orfatigue of one or more components; maximum thrust output; engineefficiency; engine noise; engine weight; engine emissions.
 6. The methodof claim 1, wherein replacing the one of the first modules comprises:replacing one or more component parts of a different fan module, enginecore module or fan case module having the same design parameters as theone of the first modules, to form the replacement module.
 7. The methodof claim 1, wherein the one of the first modules comprises the fanmodule or the engine core module, the method further comprising: passingoff the reassembled engine or the replacement module, to determine apower setting parameter of the engine or a power setting parameter orpower rating of the replacement module for use in controlling theengine, during operation.
 8. The method of claim 1, whereindisassembling the gas turbine engine comprises: removing the first fanmodule from engine core; disengaging joints between the first enginecore module and the first fan case module; and separating the enginecore from the fan case in an axial direction.
 9. The method of claim 1,including replacing another of the fan module, engine core module or fancase module with a second replacement module, prior to reassembling, thesecond replacement module being designed to different parameters to theother of the first fan module, first engine core module or first fancase module, such that it alters the performance of the engine.
 10. Themethod of claim 1, wherein the gearbox is a first gearbox, the methodfurther including: replacing the first gearbox in the engine core modulewith a replacement gearbox prior to reassembling, the replacementgearbox being designed to different parameters to the first gearbox. 11.The method of claim 1, wherein: the turbine is a first turbine, thecompressor is a first compressor, and the core shaft is a first coreshaft; the engine core further comprises a second turbine, a secondcompressor, and a second core shaft connecting the second turbine to thesecond compressor; and the second turbine, second compressor, and secondcore shaft are arranged to rotate at a higher rotational speed than thefirst core shaft.
 12. The method of claim 1, wherein the engine coremodule and fan module are passed off separately, and wherein the methodfurther comprises: passing off the replacement module independently ofthe other of the first modules after reassembly of the engine.
 13. Asystem comprising a plurality of gas turbine engines, the system having:a plurality of first engine core modules, each comprising: an enginecore having a turbine, a compressor, a core shaft connecting the turbineto the compressor; and a gearbox that receives an input from the coreshaft and outputs drive at a lower rotational speed than the core shaft;a plurality of first fan modules compatible with the engine core firstmodules, each first fan module comprising a fan arranged to receivedrive from an engine core and having a plurality of fan blades; aplurality of first fan case modules compatible with the first enginecore modules and the first fan modules, each first fan case modulecomprising a fan case arranged to surround a fan and at least partiallydefining a bypass duct around an engine core; and a plurality ofreplacement modules comprising replacements for a one of the first fanmodules, the first engine core modules or the first fan case module,wherein the replacement modules are compatible with the others of thefirst fan modules, the first engine core modules or the first fan casemodules; wherein each of the plurality of gas turbine engines comprisesa first fan module from the plurality of first fan modules, a firstengine core module from the plurality of first engine core modules and afirst fan case module from the plurality of first fan case modules; andwherein the replacement modules are interchangeable with the one of thefirst fan modules, the first engine core modules or the first fan casemodule, such that it alters the performance of the engine.
 14. Thesystem of claim 13, wherein: the one of the first modules are arrangedto be releasably secured to the others of the first modules through oneor more connection points at one or more connection positions; and thereplacement modules are arranged to be releasably secured to the othersof the first modules through the one or more connection points at theone or more connection positions.
 15. The system of claim 13, wherein:the one of the first modules includes one or more apertures, openings orelectrical connections arranged to be coupled to the others of the firstmodules in first coupling positions; and the replacement modules includethe same one or more apertures, openings or electrical connections arearranged to be coupled to the others of the first modules in the firstcoupling positions.
 16. The system of claim 13, wherein the replacementmodule is arranged to provide improved performance of the gas turbineengine after reassembly.
 17. The system of claim 16, wherein theimproved performance may include one or more of: resistance to wear andtear or fatigue of one or more components; maximum thrust output; engineefficiency; engine noise; engine weight; engine emissions.
 18. Thesystem of claim 13, wherein the engine core modules and fan modules arepassed off separately.